Turbine engine assembly and dual fuel aircraft system

ABSTRACT

A turbine engine assembly including a turbine core and a cryogenic fuel system. The turbine core includes: a compressor section; a combustion section; and a turbine section, which are axially aligned. The a cryogenic fuel system includes: a cryogenic fuel reservoir; a vaporizer heat exchanger; a liquid supply line operably coupling the fuel reservoir to an input of the vaporizer heat exchanger; a gas supply line operably coupling an output of the vaporizer heat exchanger to the combustion section; and a second heat exchanger thermally connecting the liquid supply line and the gas supply line to transfer heat from the gas supply line to the liquid supply line.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of U.S. Provisional PatentApplication No. 61/746,739, filed on Dec. 28, 2012, and PCT ApplicationNo. PCT/US2013/071794, filed Nov. 26, 2013, which are incorporatedherein in their entirety.

BACKGROUND

The technology described herein relates generally to aircraft systems,and more specifically to aircraft systems using dual fuels in anaviation gas turbine engine and a method of operating same.

Some aircraft engines may be configured to operate using one or morefuels, such as jet fuel and/or natural gas.

BRIEF DESCRIPTION OF EMBODIMENTS OF THE INVENTION

In one aspect, an embodiment of the invention relates to a turbineengine assembly, including a turbine core, having a compressor section,a combustion section, a turbine section, and a nozzle section, which areaxially aligned and a cryogenic fuel system, having a cryogenic fuelreservoir, a vaporizer heat exchanger located within the nozzle section,a liquid supply line operably coupling the fuel reservoir to an input ofthe vaporizer heat exchanger, a gas supply line operably coupling anoutput of the vaporizer heat exchanger to the combustion section, and asecond heat exchanger thermally connecting the liquid supply line andthe gas supply line to transfer heat from the gas supply line to theliquid supply line.

In another aspect, an embodiment of the invention relates to a dual fuelaircraft system for a turbine engine of an aircraft, including a firstfuel system for controlling the flow of a first fuel from a first fueltank to the turbine engine and a second fuel system for controlling theflow of cryogenic fuel to the turbine engine, having a cryogenic fuelreservoir, a vaporizer heat exchanger located within the nozzle section,a liquid supply line operably coupling the fuel reservoir to an input ofthe vaporizer heat exchanger, a gas supply line operably coupling anoutput of the vaporizer heat exchanger to the combustion section, and asecond heat exchanger thermally connecting the liquid supply line andthe gas supply line to transfer heat from the gas supply line to theliquid supply line.

BRIEF DESCRIPTION OF THE DRAWINGS

The technology described herein may be best understood by reference tothe following description taken in conjunction with the accompanyingdrawing figures in which:

FIG. 1 is an isometric view of an exemplary aircraft system having adual fuel propulsion system;

FIG. 2 is an exemplary fuel delivery/distribution system;

FIG. 2a is an exemplary operating path in a schematic pressure-enthalpychart of an exemplary cryogenic fuel;

FIG. 3 is a schematic figure showing exemplary arrangement of a fueltank and exemplary boil off usage;

FIG. 4 is a schematic cross-sectional view of an exemplary dual fuelaircraft gas turbine engine having a fuel delivery and control system;

FIG. 5 is a schematic cross-sectional view of a portion of an exemplarydual fuel aircraft gas turbine engine showing a schematic heatexchanger;

FIG. 6a is a schematic view of an exemplary direct heat exchanger;

FIG. 6b is a schematic view of an exemplary indirect heat exchanger;

FIG. 6c is a schematic view of another exemplary indirect heatexchanger;

FIG. 7 is a schematic plot of an exemplary flight mission profile forthe aircraft system; and

FIG. 8 is a schematic view of an example vaporizer system including aregenerator heat exchanger, all according to at least some aspects ofthe present disclosure.

DESCRIPTION OF EMBODIMENTS OF THE INVENTION

In the following detailed description, reference is made to theaccompanying drawings, which form a part hereof. In the drawings,similar symbols typically identify similar components, unless contextdictates otherwise. The illustrative embodiments described in thedescription, drawings, and claims are not meant to be limiting. Otherembodiments may be utilized, and other changes may be made, withoutdeparting from the spirit or scope of the subject matter presented here.It will be readily understood that the aspects of the presentdisclosure, as generally described herein, and illustrated in thefigures, can be arranged, substituted, combined, and designed in a widevariety of different configurations, all of which are explicitlycontemplated and make part of this disclosure.

FIG. 1 shows an aircraft system 5 according to an exemplary embodimentof the present invention. The exemplary aircraft system 5 has a fuselage6 and wings 7 attached to the fuselage. The aircraft system 5 has apropulsion system 100 that produces the propulsive thrust required topropel the aircraft system in flight. Although the propulsion system 100is shown attached to the wing 7 in FIG. 1, in other embodiments it maybe coupled to other parts of the aircraft system 5, such as, forexample, the tail portion 16.

The exemplary aircraft system 5 has a fuel storage system 10 for storingone or more types of fuels that are used in the propulsion system 100.The exemplary aircraft system 5 shown in FIG. 1 uses two types of fuels,as explained further below herein. Accordingly, the exemplary aircraftsystem 5 comprises a first fuel tank 21 capable of storing a first fuel11 and a second fuel tank 22 capable of storing a second fuel 12. In theexemplary aircraft system 5 shown in FIG. 1, at least a portion of thefirst fuel tank 21 is located in a wing 7 of the aircraft system 5. Inone exemplary embodiment, shown in FIG. 1, the second fuel tank 22 islocated in the fuselage 6 of the aircraft system near the location wherethe wings are coupled to the fuselage. In alternative embodiments, thesecond fuel tank 22 may be located at other suitable locations in thefuselage 6 or the wing 7. In other embodiments, the aircraft system 5may comprise an optional third fuel tank 123 capable of storing thesecond fuel 12. The optional third fuel tank 123 may be located in anaft portion of the fuselage of the aircraft system, such as for exampleshown schematically in FIG. 1.

As further described later herein, the propulsion system 100 shown inFIG. 1 is a dual fuel propulsion system that is capable of generatingpropulsive thrust by using the first fuel 11 or the second fuel 12 orusing both first fuel 11 and the second fuel 12. The exemplary dual fuelpropulsion system 100 comprises a gas turbine engine 101 capable ofgenerating a propulsive thrust selectively using the first fuel 11, orthe second fuel 21, or using both the first fuel and the second fuel atselected proportions. The first fuel may be a conventional liquid fuelsuch as a kerosene based jet fuel such as known in the art as Jet-A,JP-8, or JP-5 or other known types or grades. In the exemplaryembodiments described herein, the second fuel 12 is a cryogenic fuelthat is stored at very low temperatures. In one embodiment describedherein, the cryogenic second fuel 12 is Liquefied Natural Gas(alternatively referred to herein as “LNG”). The cryogenic second fuel12 is stored in the fuel tank at a low temperature. For example, the LNGis stored in the second fuel tank 22 at about −265° F. at an absolutepressure of about 15 psia. The fuel tanks may be made from knownmaterials such as titanium, Inconel, aluminum or composite materials.

The exemplary aircraft system 5 shown in FIG. 1 comprises a fueldelivery system 50 capable of delivering a fuel from the fuel storagesystem 10 to the propulsion system 100. Known fuel delivery systems maybe used for delivering the conventional liquid fuel, such as the firstfuel 11. In the exemplary embodiments described herein, and shown inFIGS. 1 and 2, the fuel delivery system 50 is configured to deliver acryogenic liquid fuel, such as, for example, LNG, to the propulsionsystem 100 through conduits 54 that transport the cryogenic fuel. Inorder to substantially maintain a liquid state of the cryogenic fuelduring delivery, at least a portion of the conduit 54 of the fueldelivery system 50 is insulated and configured for transporting apressurized cryogenic liquid fuel. In some exemplary embodiments, atleast a portion of the conduit 54 has a double wall construction. Theconduits may be made from known materials such as titanium, Inconel,aluminum or composite materials.

The exemplary embodiment of the aircraft system 5 shown in FIG. 1further includes a fuel cell system 400, comprising a fuel cell capableof producing electrical power using at least one of the first fuel 11 orthe second fuel 12. The fuel delivery system 50 is capable of deliveringa fuel from the fuel storage system 10 to the fuel cell system 400. Inone exemplary embodiment, the fuel cell system 400 generates power usinga portion of a cryogenic fuel 12 used by a dual fuel propulsion system100.

The propulsion system 100 comprises a gas turbine engine 101 thatgenerates the propulsive thrust by burning a fuel in a combustor. FIG. 4is a schematic view of an exemplary gas turbine engine 101 including afan 103 and a core engine 108 having a high pressure compressor 105, anda combustor 90. Engine 101 also includes a high pressure turbine 155, alow pressure turbine 157, and a booster 104. The exemplary gas turbineengine 101 has a fan 103 that produces at least a portion of thepropulsive thrust. Engine 101 has an intake side 109 and an exhaust side110. Fan 103 and turbine 157 are coupled together using a first rotorshaft 114, and compressor 105 and turbine 155 are coupled together usinga second rotor shaft 115. In some applications, such as, for example,shown in FIG. 4, the fan 103 blade assemblies are at least partiallypositioned within an engine casing 116. In other applications, the fan103 may form a portion of an “open rotor” where there is no casingsurrounding the fan blade assembly.

During operation, air flows axially through fan 103, in a direction thatis substantially parallel to a central line axis 15 extending throughengine 101, and compressed air is supplied to high pressure compressor105. The highly compressed air is delivered to combustor 90. Hot gases(not shown in FIG. 4) from combustor 90 drives turbines 155 and 157.Turbine 157 drives fan 103 by way of shaft 114 and similarly, turbine155 drives compressor 105 by way of shaft 115. In alternativeembodiments, the engine 101 may have an additional compressor, sometimesknown in the art as an intermediate pressure compressor, driven byanother turbine stage (not shown in FIG. 4).

During operation of the aircraft system 5 (See exemplary flight profileshown in FIG. 7), the gas turbine engine 101 in the propulsion system100 may use, for example, the first fuel 11 during a first selectedportion of operation of propulsion system, such as for example, duringtake off. The propulsion system 100 may use the second fuel 12, such as,for example, LNG, during a second selected portion of operation ofpropulsion system such as during cruise. Alternatively, during selectedportions of the operation of the aircraft system 5, the gas turbineengine 101 is capable of generating the propulsive thrust using both thefirst fuel 11 and the second fuel 12 simultaneously. The proportion ofthe first fuel and second fuel may be varied between 0% to 100% asappropriate during various stages of the operation of the propulsionsystem.

An aircraft and engine system, described herein, is capable of operationusing two fuels, one of which may be a cryogenic fuel such as forexample, LNG (liquefied natural gas), the other a conventional kerosenebased jet fuel such as Jet-A, JP-8, JP-5 or similar grades availableworldwide.

The Jet-A fuel system is similar to conventional aircraft fuel systems,with the exception of the fuel nozzles, which are capable of firingJet-A and cryogenic/LNG to the combustor in proportions from 0-100%. Inthe embodiment shown in FIG. 1, the LNG system includes a fuel tank,which optionally contains the following features: (i) vent lines withappropriate check valves to maintain a specified pressure in the tank;(ii) drain lines for the liquid cryogenic fuel; (iii) gauging or othermeasurement capability to assess the temperature, pressure, and volumeof cryogenic (LNG) fuel present in the tank; (iv) a boost pump locatedin the cryogenic (LNG) tank or optionally outside of the tank, whichincreases the pressure of the cryogenic (LNG) fuel to transport it tothe engine; and (iv) an optional cryo-cooler to keep the tank atcryogenic temperatures indefinitely.

The fuel tank will preferably operate at or near atmospheric pressure,but can operate in the range of 0 to 100 psig. Alternative embodimentsof the fuel system may include high tank pressures and temperatures. Thecryogenic (LNG) fuel lines running from the tank and boost pump to theengine pylons may have the following features: (i) single or double wallconstruction; (ii) vacuum insulation or low thermal conductivitymaterial insulation; and (iii) an optional cryo-cooler to re-circulateLNG flow to the tank without adding heat to the LNG tank. The cryogenic(LNG) fuel tank can be located in the aircraft where a conventionalJet-A auxiliary fuel tank is located on existing systems, for example,in the forward or aft cargo hold. Alternatively, a cryogenic (LNG) fueltank can be located in the center wing tank location. An auxiliary fueltank utilizing cryogenic (LNG) fuel may be designed so that it can beremoved if cryogenic (LNG) fuel will not be used for an extended periodof time.

A high pressure pump may be located in the pylon or on board the engineto raise the pressure of the cryogenic (LNG) fuel to levels sufficientto inject fuel into the gas turbine combustor. The pump may or may notraise the pressure of the LNG/cryogenic liquid above the criticalpressure (Pc) of cryogenic (LNG) fuel. A heat exchanger, referred toherein as a “vaporizer,” which may be mounted on or near the engine,adds thermal energy to the liquefied natural gas fuel, raising thetemperature and volumetrically expanding the cryogenic (LNG) fuel. Heat(thermal energy) from the vaporizer can come from many sources. Theseinclude, but are not limited to: (i) the gas turbine exhaust; (ii)compressor intercooling; (iii) high pressure and/or low pressure turbineclearance control air; (iv) LPT pipe cooling parasitic air; (v) cooledcooling air from the HP turbine; (vi) lubricating oil; or (vii) on boardavionics or electronics. The heat exchanger can be of various designs,including shell and tube, double pipe, fin plate, etc., and can flow ina co-current, counter current, or cross current manner. Heat exchangecan occur in direct or indirect contact with the heat sources listedabove.

A control valve is located downstream of the vaporizer/heat exchangeunit described above. The purpose of the control valve is to meter theflow to a specified level into the fuel manifold across the range ofoperational conditions associated with the gas turbine engine operation.A secondary purpose of the control valve is to act as a back pressureregulator, setting the pressure of the system above the criticalpressure of cryogenic (LNG) fuel.

A fuel manifold is located downstream of the control valve, which servesto uniformly distribute gaseous fuel to the gas turbine fuel nozzles. Insome embodiments, the manifold can optionally act as a heat exchanger,transferring thermal energy from the core cowl compartment or otherthermal surroundings to the cryogenic/LNG/natural gas fuel. A purgemanifold system can optionally be employed with the fuel manifold topurge the fuel manifold with compressor air (CDP) when the gaseous fuelsystem is not in operation. This will prevent hot gas ingestion into thegaseous fuel nozzles due to circumferential pressure variations.Optionally, check valves in or near the fuel nozzles can prevent hot gasingestion.

An exemplary embodiment of the system described herein may operate asfollows: Cryogenic (LNG) fuel is located in the tank at about 15 psiaand about −265° F. It is pumped to approximately 30 psi by the boostpump located on the aircraft. Liquid cryogenic (LNG) fuel flows acrossthe wing via insulated double walled piping to the aircraft pylon whereit is stepped up to about 100 to 1,500 psia and can be above or belowthe critical pressure of natural gas/methane. The cryogenic (LNG) fuelis then routed to the vaporizer where it volumetrically expands to agas. The vaporizer may be sized to keep the Mach number andcorresponding pressure losses low. Gaseous natural gas is then meteredthough a control valve and into the fuel manifold and fuel nozzles whereit is combusted in an otherwise standard aviation gas turbine enginesystem, providing thrust to the airplane. As cycle conditions change,the pressure in the boost pump (about 30 psi for example) and thepressure in the HP pump (about 1,000 psi for example) are maintained atan approximately constant level. Flow is controlled by the meteringvalve. The variation in flow in combination with the appropriately sizedfuel nozzles result in acceptable and varying pressures in the manifold.

The exemplary aircraft system 5 has a fuel delivery system fordelivering one or more types of fuels from the storage system 10 for usein the propulsion system 100. For a conventional liquid fuel such as,for example, a kerosene based jet fuel, a conventional fuel deliverysystem may be used. The exemplary fuel delivery system described herein,and shown schematically in FIGS. 2 and 3, comprises a cryogenic fueldelivery system 50 for an aircraft system 5. The exemplary fuel system50 shown in FIG. 2 comprises a cryogenic fuel tank 122 capable ofstoring a cryogenic liquid fuel 112. In one embodiment, the cryogenicliquid fuel 112 is LNG. Other alternative cryogenic liquid fuels mayalso be used. In the exemplary fuel system 50, the cryogenic liquid fuel112, such as, for example, LNG, is at a first pressure “P1”. Thepressure P1 is preferably close to atmospheric pressure, such as, forexample, 15 psia.

The exemplary fuel system 50 has a boost pump 52 such that it is in flowcommunication with the cryogenic fuel tank 122. During operation, whencryogenic fuel is needed in the dual fuel propulsion system 100, theboost pump 52 removes a portion of the cryogenic liquid fuel 112 fromthe cryogenic fuel tank 122 and increases its pressure to a secondpressure “P2” and flows it into a wing supply conduit 54 located in awing 7 of the aircraft system 5. The pressure P2 is chosen such that theliquid cryogenic fuel maintains its liquid state (L) during the flow inthe supply conduit 54. The pressure P2 may be in the range of about 30psia to about 40 psia. Based on analysis using known methods, for LNG,30 psia is found to be adequate. The boost pump 52 may be located at asuitable location in the fuselage 6 of the aircraft system 5.Alternatively, the boost pump 52 may be located close to the cryogenicfuel tank 122. In other embodiments, the boost pump 52 may be locatedinside the cryogenic fuel tank 122. In order to substantially maintain aliquid state of the cryogenic fuel during delivery, at least a portionof the wing supply conduit 54 is insulated. In some exemplaryembodiments, at least a portion of the conduit 54 has a double wallconstruction. The conduits 54 and the boost pump 52 may be made usingknown materials such as titanium, Inconel, aluminum or compositematerials.

The exemplary fuel system 50 has a high-pressure pump 58 that is in flowcommunication with the wing supply conduit 54 and is capable ofreceiving the cryogenic liquid fuel 112 supplied by the boost pump 52.The high-pressure pump 58 increases the pressure of the liquid cryogenicfuel (such as, for example, LNG) to a third pressure “P3” sufficient toinject the fuel into the propulsion system 100. The pressure P3 may bein the range of about 100 psia to about 1000 psia. The high-pressurepump 58 may be located at a suitable location in the aircraft system 5or the propulsion system 100. The high-pressure pump 58 is preferablylocated in a pylon 55 of aircraft system 5 that supports the propulsionsystem 100.

As shown in FIG. 2, the exemplary fuel system 50 has a vaporizer 60 forchanging the cryogenic liquid fuel 112 into a gaseous (G) fuel 13. Thevaporizer 60 receives the high pressure cryogenic liquid fuel and addsheat (thermal energy) to the cryogenic liquid fuel (such as, forexample, LNG) raising its temperature and volumetrically expanding it.Heat (thermal energy) can be supplied from one or more sources in thepropulsion system 100. For example, heat for vaporizing the cryogenicliquid fuel in the vaporizer may be supplied from one or more of severalsources, such as, for example, the gas turbine exhaust 99, compressor105, high pressure turbine 155, low pressure turbine 157, fan bypass107, turbine cooling air, lubricating oil in the engine, aircraft systemavionics/electronics, or any source of heat in the propulsion system100. Due to the exchange of heat that occurs in the vaporizer 60, thevaporizer 60 may be alternatively referred to as a heat exchanger. Theheat exchanger portion of the vaporizer 60 may include a shell and tubetype heat exchanger, or a double pipe type heat exchanger, orfin-and-plate type heat exchanger. The hot fluid and cold fluid flow inthe vaporizer may be co-current, or counter-current, or a cross currentflow type. The heat exchange between the hot fluid and the cold fluid inthe vaporizer may occur directly through a wall or indirectly, using anintermediate work fluid.

The cryogenic fuel delivery system 50 comprises a flow metering valve 65(“FMV”, also referred to as a Control Valve) that is in flowcommunication with the vaporizer 60 and a manifold 70. The flow meteringvalve 65 is located downstream of the vaporizer/heat exchange unitdescribed above. The purpose of the FMV (control valve) is to meter thefuel flow to a specified level into the fuel manifold 70 across therange of operational conditions associated with the gas turbine engineoperation. A secondary purpose of the control valve is to act as a backpressure regulator, setting the pressure of the system above thecritical pressure of the cryogenic fuel such as LNG. The flow meteringvalve 65 receives the gaseous fuel 13 supplied from the vaporizer andreduces its pressure to a fourth pressure “P4”. The manifold 70 iscapable of receiving the gaseous fuel 13 and distributing it to a fuelnozzle 80 in the gas turbine engine 101. In a preferred embodiment, thevaporizer 60 changes the cryogenic liquid fuel 112 into the gaseous fuel13 at a substantially constant pressure. FIG. 2a schematically shows thestate and pressure of the fuel at various points in the delivery system50.

The cryogenic fuel delivery system 50 further comprises a plurality offuel nozzles 80 located in the gas turbine engine 101. The fuel nozzle80 delivers the gaseous fuel 13 into the combustor 90 for combustion.The fuel manifold 70, located downstream of the control valve 65, servesto uniformly distribute gaseous fuel 13 to the gas turbine fuel nozzles80. In some embodiments, the manifold 70 can optionally act as a heatexchanger, transferring thermal energy from the propulsion system corecowl compartment or other thermal surroundings to the LNG/natural gasfuel. In one embodiment, the fuel nozzle 80 is configured to selectivelyreceive a conventional liquid fuel (such as the conventional kerosenebased liquid fuel) or the gaseous fuel 13 generated by the vaporizerfrom the cryogenic liquid fuel such as LNG. In another embodiment, thefuel nozzle 80 is configured to selectively receive a liquid fuel andthe gaseous fuel 13 and configured to supply the gaseous fuel 13 and aliquid fuel to the combustor 90 to facilitate co-combustion of the twotypes of fuels. In another embodiment, the gas turbine engine 101comprises a plurality of fuel nozzles 80 wherein some of the fuelnozzles 80 are configured to receive a liquid fuel and some of the fuelnozzles 80 are configured to receive the gaseous fuel 13 and arrangedsuitably for combustion in the combustor 90.

In another embodiment of the present invention, fuel manifold 70 in thegas turbine engine 101 comprises an optional purge manifold system topurge the fuel manifold with compressor air, or other air, from theengine when the gaseous fuel system is not in operation. This willprevent hot gas ingestion into the gaseous fuel nozzles due tocircumferential pressure variations in the combustor 90. Optionally,check valves in or near the fuel nozzles can be used prevent hot gasingestion in the fuel nozzles or manifold.

In an exemplary dual fuel gas turbine propulsion system described hereinthat uses LNG as the cryogenic liquid fuel is described as follows: LNGis located in the tank 22, 122 at 15 psia and −265° F. It is pumped toapproximately 30 psi by the boost pump 52 located on the aircraft.Liquid LNG flows across the wing 7 via insulated double walled piping 54to the aircraft pylon 55 where it is stepped up to 100 to 1,500 psia andmay be above or below the critical pressure of natural gas/methane. TheLiquefied Natural Gas is then routed to the vaporizer 60 where itvolumetrically expands to a gas. The vaporizer 60 is sized to keep theMach number and corresponding pressure losses low. Gaseous natural gasis then metered though a control valve 65 and into the fuel manifold 70and fuel nozzles 80 where it is combusted in an dual fuel aviation gasturbine system 100, 101, providing thrust to the aircraft system 5. Ascycle conditions change, the pressure in the boost pump (30 psi) and thepressure in the HP pump 58 (1,000 psi) are maintained at anapproximately constant level. Flow is controlled by the metering valve65. The variation in flow in combination with the appropriately sizedfuel nozzles result in acceptable and varying pressures in the manifold.

The dual fuel system consists of parallel fuel delivery systems forkerosene based fuel (Jet-A, JP-8, JP-5, etc) and a cryogenic fuel (LNGfor example). The kerosene fuel delivery is substantially unchanged fromthe current design, with the exception of the combustor fuel nozzles,which are designed to co-fire kerosene and natural gas in anyproportion. As shown in FIG. 2, the cryogenic fuel (LNG for example)fuel delivery system consists of the following features: (A) A dual fuelnozzle and combustion system, capable of utilizing cryogenic fuel (LNGfor example), and Jet-A in any proportion from 0- to 100%; (B) A fuelmanifold and delivery system that also acts as a heat exchanger, heatingcryogenic fuel (LNG for example) to a gas or a supercritical fluid. Themanifold system is designed to concurrently deliver fuel to thecombustor fuel nozzles in a uniform manner, and absorb heat from thesurrounding core cowl, exhaust system, or other heat source, eliminatingor minimizing the need for a separate heat exchanger; (C) A fuel systemthat pumps up cryogenic fuel (LNG for example) in its liquid state aboveor below the critical pressure and adds heat from any of a number ofsources; (D) A low pressure cryo-pump submerged in the cryogenic fuel(LNG for example) fuel tank (optionally located outside the fuel tank.);(E) A high pressure cryo-pump located in the aircraft pylon oroptionally on board the engine or nacelle to pump to pressures above thecritical pressure of cryogenic fuel (LNG for example). (F) A purgemanifold system can optionally employed with the fuel manifold to purgethe fuel manifold with compressor CDP air when the gaseous fuel systemis not in operation. This will prevent hot gas ingestion into thegaseous fuel nozzles due to circumferential pressure variations.Optionally, check valves in or near the fuel nozzles can prevent hot gasingestion. (G) cryogenic fuel (LNG for example) lines running from thetank and boost pump to the engine pylons have the following features:(1) Single or double wall construction. (2) Vacuum insulation oroptionally low thermal conductivity insulation material such asaerogels. (3) An optional cryo-cooler to recirculate cryogenic fuel (LNGfor example) flow to the tank without adding heat to the cryogenic fuel(LNG for example) tank. (H) A high pressure pump located in the pylon oron board the engine. This pump will raise the pressure of the cryogenicfuel (LNG for example) to levels sufficient to inject natural gas fuelinto the gas turbine combustor. The pump may or may not raise thepressure of the cryogenic liquid (LNG for example) above the criticalpressure (Pc) of cryogenic fuel (LNG for example).

III. A Fuel Storage System

The exemplary aircraft system 5 shown in FIG. 1 comprises a cryogenicfuel storage system 10, such as shown for example, in FIG. 3, forstoring a cryogenic fuel. The exemplary cryogenic fuel storage system 10comprises a cryogenic fuel tank 22, 122 having a first wall 23 forming astorage volume 24 capable of storing a cryogenic liquid fuel 12 such asfor example LNG. As shown schematically in FIG. 3, the exemplarycryogenic fuel storage system 10 has an inflow system 32 capable offlowing the cryogenic liquid fuel 12 into the storage volume 24 and anoutflow system 30 adapted to deliver the cryogenic liquid fuel 12 fromthe cryogenic fuel storage system 10. It further comprises a vent system40 capable of removing at least a portion of a gaseous fuel 19 (that maybe formed during storage) from the cryogenic liquid fuel 12 in thestorage volume 24.

The exemplary cryogenic fuel storage system 10 shown in FIG. 3 furthercomprises a recycle system 34 that is adapted to return at least aportion 29 of unused gaseous fuel 19 into the cryogenic fuel tank 22. Inone embodiment, the recycle system 34 comprises a cryo-cooler 42 thatcools the portion 29 of unused gaseous fuel 19 prior to returning itinto the cryogenic fuel tank 22, 122. An exemplary operation of thecryo-cooler 42 operation is as follows: In an exemplary embodiment, boiloff from the fuel tank can be re-cooled using a reverse Rankinerefrigeration system, also known as a cryo-cooler. The cryo-cooler canbe powered by electric power coming from any of the available systems onboard the aircraft system 5, or, by ground based power systems such asthose which may be available while parked at a boarding gate. Thecryo-cooler system can also be used to re-liquefy natural gas in thefuel system during the dual fuel aircraft gas turbine engine 101 co-firetransitions.

The fuel storage system 10 may further comprise a safety release system45 adapted to vent any high pressure gases that may be formed in thecryogenic fuel tank 22. In one exemplary embodiment, shown schematicallyin FIG. 3, the safety release system 45 comprises a rupture disk 46 thatforms a portion of the first wall 23. The rupture disk 46 is a safetyfeature, designed using known methods, to blow out and release any highpressure gases in the event of an over pressure inside the fuel tank 22.

The cryogenic fuel tank 22 may have a single wall construction or amultiple wall construction. For example, the cryogenic fuel tank 22 mayfurther comprise (See FIG. 3 for example) a second wall 25 thatsubstantially encloses the first wall 23. In one embodiment of the tank,there is a gap 26 between the first wall 23 and the second wall 25 inorder to thermally insulate the tank to reduce heat flow across the tankwalls. In one exemplary embodiment, there is a vacuum in the gap 26between the first wall 23 and the second wall 25. The vacuum may becreated and maintained by a vacuum pump 28. Alternatively, in order toprovide thermal insulation for the tank, the gap 26 between the firstwall 23 and the second wall 25 may be substantially filled with a knownthermal insulation material 27, such as, for example, Aerogel. Othersuitable thermal insulation materials may be used. Baffles 17 may beincluded to control movement of liquid within the tank.

The cryogenic fuel storage system 10 shown in FIG. 3 comprises theoutflow system 30 having a delivery pump 31. The delivery pump may belocated at a convenient location near the tank 22. In order to reduceheat transfer in to the cryogenic fuel, it may be preferable to locatethe delivery pump 31 in the cryogenic fuel tank 22 as shownschematically in FIG. 3. The vent system 40 vents any gases that may beformed in the fuel tank 22. These vented gases may be utilized inseveral useful ways in the aircraft system 5. A few of these are shownschematically in FIG. 3. For example at least a portion of the gaseousfuel 19 may be supplied to the aircraft propulsion system 100 forcooling or combustion in the engine. In another embodiment, the ventsystem 40 supplies at least a portion of the gaseous fuel 19 to a burnerand further venting the combustion products from the burner safelyoutside the aircraft system 5. In another embodiment the vent system 40supplies at least a portion of the gaseous fuel 19 to an auxiliary powerunit 180 that supplies auxiliary power to the aircraft system 5. Inanother embodiment the vent system 40 supplies at least a portion of thegaseous fuel 19 to a fuel cell 182 that produces power. In anotherembodiment the vent system 40 releases at least a portion of the gaseousfuel 19 outside the cryogenic fuel tank 22.

The exemplary operation of the fuel storage system, its componentsincluding the fuel tank, and exemplary sub systems and components isdescribed as follows.

Natural gas exists in liquid form (LNG) at temperatures of approximatelyabout −260° F. and atmospheric pressure. To maintain these temperaturesand pressures on board a passenger, cargo, military, or general aviationaircraft, the features identified below, in selected combinations, allowfor safe, efficient, and cost effective storage of LNG. Referring toFIG. 3, these include:

A fuel tank 21, 22 constructed of alloys such as, but not limited to,aluminum AL 5456 and higher strength aluminum AL 5086 or other suitablealloys.

A fuel tank 21, 22 constructed of light weight composite material.

The above tanks 21, 22 with a double wall vacuum feature for improvedinsulation and greatly reduced heat flow to the LNG fluid. The doublewalled tank also acts as a safety containment device in the rare casewhere the primary tank is ruptured.

An alternative embodiment of either the above utilizing lightweightinsulation 27, such as, for example, Aerogel, to minimize heat flow fromthe surroundings to the LNG tank and its contents. Aerogel insulationcan be used in addition to, or in place of a double walled tank design.

An optional vacuum pump 28 designed for active evacuation of the spacebetween the double walled tank. The pump can operate off of LNG boil offfuel, LNG, Jet-A, electric power or any other power source available tothe aircraft.

An LNG tank with a cryogenic pump 31 submerged inside the primary tankfor reduced heat transfer to the LNG fluid.

An LNG tank with one or more drain lines 36 capable of removing LNG fromthe tank under normal or emergency conditions. The LNG drain line 36 isconnected to a suitable cryogenic pump to increase the rate of removalbeyond the drainage rate due to the LNG gravitational head.

An LNG tank with one or more vent lines 41 for removal of gaseousnatural gas, formed by the absorption of heat from the externalenvironment. This vent line 41 system maintains the tank at a desiredpressure by the use of a 1 way relief valve or back pressure valve 39.

An LNG tank with a parallel safety relief system 45 to the main ventline, should an overpressure situation occur. A burst disk is analternative feature or a parallel feature 46. The relief vent woulddirect gaseous fuel overboard.

An LNG fuel tank, with some or all of the design features above, whosegeometry is designed to conform to the existing envelope associated witha standard Jet-A auxiliary fuel tank such as those designed andavailable on commercially available aircrafts.

An LNG fuel tank, with some or all of the design features above, whosegeometry is designed to conform to and fit within the lower cargohold(s) of conventional passenger and cargo aircraft such as those foundon commercially available aircrafts.

Modifications to the center wing tank 22 of an existing or new aircraftto properly insulate the LNG, tank, and structural elements.

Venting and boil off systems are designed using known methods. Boil offof LNG is an evaporation process which absorbs energy and cools the tankand its contents. Boil off LNG can be utilized and/or consumed by avariety of different processes, in some cases providing useful work tothe aircraft system, in other cases, simply combusting the fuel for amore environmentally acceptable design. For example, vent gas from theLNG tank consists primarily of methane and is used for any or allcombinations of the following:

Routing to the Aircraft APU (Auxiliary Power Unit) 180. As shown in FIG.3, a gaseous vent line from the tank is routed in series or in parallelto an Auxiliary Power Unit for use in the combustor. The APU can be anexisting APU, typically found aboard commercial and military aircraft,or a separate APU dedicated to converting natural gas boil off to usefulelectric and/or mechanical power. A boil off natural gas compressor isutilized to compress the natural gas to the appropriate pressurerequired for utilization in the APU. The APU, in turn, provides electricpower to any system on the engine or A/C.

Routing to one or more aircraft gas turbine engine(s) 101. As shown inFIG. 3, a natural gas vent line from the LNG fuel tank is routed to oneor more of the main gas turbine engines 101 and provides an additionalfuel source to the engine during operation. A natural gas compressor isutilized to pump the vent gas to the appropriate pressure required forutilization in the aircraft gas turbine engine.

Flared. As shown in FIG. 3, a natural gas vent line from the tank isrouted to a small, dedicated vent combustor 190 with its own electricspark ignition system. In this manner methane gas is not released to theatmosphere. The products of combustion are vented, which results in amore environmentally acceptable system.

Vented. As shown in FIG. 3, a natural gas vent line from the tank isrouted to the exhaust duct of one or more of the aircraft gas turbines.Alternatively, the vent line can be routed to the APU exhaust duct or aseparate dedicated line to any of the aircraft trailing edges. Naturalgas may be suitably vented to atmosphere at one or more of theselocations V.

Ground operation. As shown in FIG. 3, during ground operation, any ofthe systems can be designed such that a vent line 41 is attached toground support equipment, which collects and utilizes the natural gasboil off in any ground based system. Venting can also take place duringrefueling operations with ground support equipment that cansimultaneously inject fuel into the aircraft LNG tank using an inflowsystem 32 and capture and reuse vent gases (simultaneous venting andfueling indicated as (S) in FIG. 3).

IV. Propulsion (Engine) System

FIG. 4 shows an exemplary dual fuel propulsion system 100 comprising agas turbine engine 101 capable of generating a propulsive thrust using acryogenic liquid fuel 112. The gas turbine engine 101 comprises acompressor 105 driven by a high-pressure turbine 155 and a combustor 90that burns a fuel and generates hot gases that drive the high-pressureturbine 155. The combustor 90 is capable of burning a conventionalliquid fuel such as kerosene based fuel. The combustor 90 is alsocapable of burning a cryogenic fuel, such as, for example, LNG, that hasbeen suitably prepared for combustion, such as, for example, by avaporizer 60. FIG. 4 shows schematically a vaporizer 60 capable ofchanging the cryogenic liquid fuel 112 into a gaseous fuel 13. The dualfuel propulsion system 100 gas turbine engine 101 further comprises afuel nozzle 80 that supplies the gaseous fuel 13 to the combustor 90 forignition. In one exemplary embodiment, the cryogenic liquid fuel 112used is Liquefied Natural Gas (LNG). In a turbo-fan type dual fuelpropulsion system 100 (shown in FIG. 4 for example) the gas turbineengine 101 comprises a fan 103 located axially forward from thehigh-pressure compressor 105. A booster 104 (shown in FIG. 4) may belocated axially between the fan 103 and the high-pressure compressor 105wherein the fan and booster are driven by a low-pressure turbine 157. Inother embodiments, the dual fuel propulsion system 100 gas turbineengine 101 may include an intermediate pressure compressor driven by anintermediate pressure turbine (both not shown in FIG. 4). The booster104 (or an intermediate pressure compressor) increases the pressure ofthe air that enters the compressor 105 and facilitates the generation ofhigher pressure ratios by the compressor 105. In the exemplaryembodiment shown in FIG. 4, the fan and the booster are driven by thelow pressure turbine 157, and the high pressure compressor is driven thehigh pressure turbine 155.

The vaporizer 60, shown schematically in FIG. 4, is mounted on or nearthe engine 101. One of the functions of the vaporizer 60 is to addthermal energy to the cryogenic fuel, such as the liquefied natural gas(LNG) fuel, raising its temperature. In this context, the vaporizerfunctions as heat exchanger. Another, function of the vaporizer 60 is tovolumetrically expand the cryogenic fuel, such as the liquefied naturalgas (LNG) fuel to a gaseous form for later combustion. Heat (thermalenergy) for use in the vaporizer 60 can come from or more of manysources in the propulsion system 100 and aircraft system 5. Theseinclude, but are not limited to: (i) The gas turbine exhaust, (ii)Compressor intercooling, (iii) High pressure and/or low pressure turbineclearance control air, (iv) LPT pipe cooling parasitic air, (v) coolingair used in the High pressure and/or low pressure turbine, (vi)Lubricating oil, and (vii) On board avionics, electronics in theaircraft system 5. The heat for the vaporizer may also be supplied fromthe compressor 105, booster 104, intermediate pressure compressor (notshown) and/or the fan bypass air stream 107 (See FIG. 4). An exemplaryembodiment using a portion of the discharge air from the compressor 105is shown in FIG. 5. A portion of the compressor discharge air 2 is bledout to the vaporizer 60, as shown by item 3 in FIG. 5. The cryogenicliquid fuel 21, such as for example, LNG, enters vaporizer 60 whereinthe heat from the airflow stream 3 is transferred to the cryogenicliquid fuel 21. In one exemplary embodiment, the heated cryogenic fuelis further expanded, as described previously herein, producing gaseousfuel 13 in the vaporizer 60. The gaseous fuel 13 is then introduced intocombustor 90 using a fuel nozzle 80 (See FIG. 5). The cooled airflow 4that exits from the vaporizer can be used for cooling other enginecomponents, such as the combustor 90 structures and/or the high-pressureturbine 155 structures. The heat exchanger portion in the vaporizer 60can be of a known design, such as for example, shell and tube design,double pipe design, and/or fin plate design. The fuel 112 flow directionand the heating fluid 96 direction in the vaporizer 60 (see FIG. 4) maybe in a co-current direction, counter-current direction, or they mayflow in a cross-current manner to promote efficient heat exchangebetween the cryogenic fuel and the heating fluid.

Heat exchange in the vaporizer 60 can occur in direct manner between thecryogenic fuel and the heating fluid, through a metallic wall. FIG. 5shows schematically a direct heat exchanger in the vaporizer 60. FIG. 6ashows schematically an exemplary direct heat exchanger 63 that uses aportion 97 of the gas turbine engine 101 exhaust gas 99 to heat thecryogenic liquid fuel 112. Alternatively, heat exchange in the vaporizer60 can occur in an indirect manner between the cryogenic fuel and theheat sources listed above, through the use of an intermediate heatingfluid. FIG. 6b shows an exemplary vaporizer 60 that uses an indirectheat exchanger 64 that uses an intermediary heating fluid 68 to heat thecryogenic liquid fuel 112. In such an indirect heat exchanger shown inFIG. 6b , the intermediary heating fluid 68 is heated by a portion 97 ofthe exhaust gas 99 from the gas turbine engine 101. Heat from theintermediary heating fluid 68 is then transferred to the cryogenicliquid fuel 112. FIG. 6c shows another embodiment of an indirectexchanger used in a vaporizer 60. In this alternative embodiment, theintermediary heating fluid 68 is heated by a portion of a fan bypassstream 107 of the gas turbine engine 101, as well as a portion 97 of theengine exhaust gas 99. The intermediary heating fluid 68 then heats thecryogenic liquid fuel 112. A control valve 38 is used to control therelative heat exchanges between the flow streams.

(V) Method of Operating Dual Fuel Aircraft System

An exemplary method of operation of the aircraft system 5 using a dualfuel propulsion system 100 is described as follows with respect to anexemplary flight mission profile shown schematically in FIG. 7. Theexemplary flight mission profile shown schematically in FIG. 7 shows theEngine power setting during various portions of the flight missionidentified by the letter labels A-B-C-D-E- . . . -X-Y etc. For example,A-B represents the start, B-C shows ground-idle, G-H shows take-off, T-Land O-P show cruise, etc. During operation of the aircraft system 5 (Seeexemplary flight profile 120 in FIG. 7), the gas turbine engine 101 inthe propulsion system 100 may use, for example, the first fuel 11 duringa first selected portion of operation of propulsion system, such as forexample, during take off. The propulsion system 100 may use the secondfuel 12, such as, for example, LNG, during a second selected portion ofoperation of propulsion system such as during cruise. Alternatively,during selected portions of the operation of the aircraft system 5, thegas turbine engine 101 is capable of generating the propulsive thrustusing both the first fuel 11 and the second fuel 12 simultaneously. Theproportion of the first fuel and second fuel may be varied between 0% to100% as appropriate during various stages of the operation of the dualfuel propulsion system 100.

An exemplary method of operating a dual fuel propulsion system 100 usinga dual fuel gas turbine engine 101 comprises the following steps of:starting the aircraft engine 101 (see A-B in FIG. 7) by burning a firstfuel 11 in a combustor 90 that generates hot gases that drive a gasturbine in the engine 101. The first fuel 11 may be a known type ofliquid fuel, such as a kerosene based Jet Fuel. The engine 101, whenstarted, may produce enough hot gases that may be used to vaporize asecond fuel, such as, for example, a cryogenic fuel. A second fuel 12 isthen vaporized using heat in a vaporizer 60 to form a gaseous fuel 13.The second fuel may be a cryogenic liquid fuel 112, such as, forexample, LNG. The operation of an exemplary vaporizer 60 has beendescribed herein previously. The gaseous fuel 13 is then introduced intothe combustor 90 of the engine 101 using a fuel nozzle 80 and thegaseous fuel 13 is burned in the combustor 90 that generates hot gasesthat drive the gas turbine in the engine. The amount of the second fuelintroduced into the combustor may be controlled using a flow meteringvalve 65. The exemplary method may further comprise the step of stoppingthe supply of the first fuel 11 after starting the aircraft engine, ifdesired.

In the exemplary method of operating the dual fuel aircraft gas turbineengine 101, the step of vaporizing the second fuel 12 may be performedusing heat from a hot gas extracted from a heat source in the engine101. As described previously, in one embodiment of the method, the hotgas may be compressed air from a compressor 155 in the engine (forexample, as shown in FIG. 5). In another embodiment of the method, thehot gas is supplied from an exhaust nozzle 98 or exhaust stream 99 ofthe engine (for example, as shown in FIG. 6a ).

The exemplary method of operating a dual fuel aircraft engine 101, may,optionally, comprise the steps of using a selected proportion of thefirst fuel 11 and a second fuel 12 during selected portions of a flightprofile 120, such as shown, for example, in FIG. 7, to generate hotgases that drive a gas turbine engine 101. The second fuel 12 may be acryogenic liquid fuel 112, such as, for example, Liquefied Natural Gas(LNG). In the method above, the step of varying the proportion of thefirst fuel 12 and the second fuel 13 during different portions of theflight profile 120 (see FIG. 7) may be used to operate the aircraftsystem in an economic and efficient manner. This is possible, forexample, in situations where the cost of the second fuel 12 is lowerthan the cost of the first fuel 11. This may be the case, for example,while using LNG as the second fuel 12 and kerosene based liquid fuelssuch as Jet-A fuel, as first fuel 11. In the exemplary method ofoperating a dual fuel aircraft engine 101, the proportion (ratio) ofamount of the second fuel 12 used to the amount of the first fuel usedmay be varied between about 0% and 100%, depending on the portion of theflight mission. For example, in one exemplary method, the proportion ofa cheaper second fuel used (such as LNG) to the kerosene based fuel usedis about 100% during a cruise part of the flight profile, in order tominimize the cost of fuel. In another exemplary operating method, theproportion of the second fuel is about 50% during a take-off part of theflight profile that requires a much higher thrust level.

The exemplary method of operating a dual fuel aircraft engine 101described above may further comprise the step of controlling the amountsof the first fuel 11 and the second fuel 12 introduced into thecombustor 90 using a control system 130. An exemplary control system 130is shown schematically in FIG. 4. The control system 130 sends a controlsignal 131 (S1) to a control valve 135 to control the amount of thefirst fuel 11 that is introduced to the combustor 90. The control system130 also sends another control signal 132 (S2) to a control valve 65 tocontrol the amount of the second fuel 12 that is introduced to thecombustor 90. The proportion of the first fuel 11 and second fuel 12used can be varied between 0% to 100% by a controller 134 that isprogrammed to vary the proportion as required during different flightsegments of the flight profile 120. The control system 130 may alsoreceive a feed back signal 133, based for example on the fan speed orthe compressor speed or other suitable engine operating parameters. Inone exemplary method, the control system may be a part of the enginecontrol system, such as, for example, a Full Authority DigitalElectronic Control (FADEC) 357. In another exemplary method, amechanical or hydromechanical engine control system may form part or allof the control system.

The control system 130, 357 architecture and strategy is suitablydesigned to accomplish economic operation of the aircraft system 5.Control system feedback to the boost pump 52 and high pressure pump(s)58 can be accomplished via the Engine FADEC 357 or by distributedcomputing with a separate control system that may, optionally,communicate with the Engine FADEC and with the aircraft system 5 controlsystem through various available data busses.

The control system, such as for example, shown in FIG. 4, item 130, mayvary pump 52, 58 speed and output to maintain a specified pressureacross the wing 7 for safety purposes (for example at about 30-40 psi)and a different pressure downstream of the high pressure pump 58 (forexample at about 100 to 1500 psi) to maintain a system pressure abovethe critical point of LNG and avoid two phase flow, and, to reduce thevolume and weight of the LNG fuel delivery system by operation at highpressures and fuel densities.

In an exemplary control system 130, 357, the control system software mayinclude any or all of the following logic: (A) A control system strategythat maximizes the use of the cryogenic fuel such as, for example, LNG,on takeoff and/or other points in the envelope at high compressordischarge temperatures (T3) and/or turbine inlet temperatures (T41); (B)A control system strategy that maximizes the use of cryogenic fuel suchas, for example, LNG, on a mission to minimize fuel costs; (C) A controlsystem 130, 357 that re-lights on the first fuel, such as, for example,Jet-A, only for altitude relights; (D) A control system 130, 357 thatperforms ground starts on conventional Jet-A only as a default setting;(E) A control system 130, 357 that defaults to Jet-A only during any nontypical maneuver; (F) A control system 130, 357 that allows for manual(pilot commanded) selection of conventional fuel (like Jet-A) orcryogenic fuel such as, for example, LNG, in any proportion; (G) Acontrol system 130, 357 that utilizes 100% conventional fuel (likeJet-A) for all fast accels and decels.

The present disclosure contemplates that the temperature of fuels thatare to be vaporized may be controlled and/or brought to a desiredtemperature for supply to nozzles for combustion in airplane engines.Some example embodiments according to at least some aspects of thepresent disclosure may aid in controlling a temperature of a fluid(e.g., a fuel, such as liquid natural gas) provided to a combustionsystem within the temperature requirements desired for combustion and/orfor the system in airplane engines. Example embodiments according to atleast some aspects of the present disclosure may be used in connectionwith various types of aircraft engines (e.g., turbo-fan, turbo-jet,turbo-prop, open-rotor, etc.). For example, an embodiment of theinvention may include a regenerator heat exchanger, which may aid withfluid temperature control in single and dual fuel engines. Theregenerator heat exchanger may be configured to transfer heat to arelatively cold fluid (with or without phase change, liquid to gas) froma relatively warm fluid. Some example warm fluids may be exiting avaporizer/heat exchanger in the exhaust gases of an airplane engine.Example heat exchanger designs include, but are not limited to, coiled,axial, and/or combinations (coiled and axial) of tubes, tube and shellheat exchanger, and/or compact, plate heat exchangers. Some exampleregenerator heat exchanger designs may be made from metal, composites,and/or a combination. Some example embodiments may be configured to varyflow through the regenerator heat exchanger and/or a bypass around theregenerator heat exchanger. By adjusting a valve in series with theregenerator heat exchanger, temperatures associated with the regeneratorheat exchanger may be controlled.

As previously described with respect to FIG. 4, a turbine engineassembly 101 may have a turbine core 108 including a compressor section105, a combustion section 90, a turbine section 155, 157 and a nozzlesection with nozzles 80, which are axially aligned. The turbine engineassembly 101 may also include a cryogenic fuel system 500, illustratedin FIG. 8. The cryogenic fuel system 500 has been illustrated asincluding a supply system including a fuel tank or cryogenic fuelreservoir 502, a vaporizer heat exchanger 504, which may be locatedwithin the nozzle section of the turbine engine assembly 101. A liquidsupply line 506 may operably couple the cryogenic fuel reservoir 502 toan input of the vaporizer heat exchanger 504. The supply system fromwhich the liquid fuel may be received may also include valving and oneor more pumps (not shown). A gas supply line 508 may operably couple anoutput of the vaporizer heat exchanger 504 to the combustion section 90of the turbine engine 101 as indicated at 509. A second heat exchanger510 may thermally connect the liquid supply line 506 and the gas supplyline 508 to transfer heat from the gas supply line 508 to the liquidsupply line 506. It will be understood that the cryogenic fuel may beLNG.

The second heat exchanger 510 may be a regenerator heat exchanger havinga first side 512 and a second side 514. As illustrated, the first side510 is fluidly coupled between the cryogenic fuel reservoir 502 and aninput to the vaporizer heat exchanger 504 and the second side 514 isfluidly coupled between the output of the vaporizer heat exchanger 504and the combustion section of the turbine engine as indicated at 509.

A liquid metering valve 516 may be in fluid communication with theliquid supply line 506 and may control the flow rate of liquid fuel.Further, a bypass line 518 having an input 520 fluidly located betweenthe output of the vaporizer heat exchanger 504 and the second side 514of the regenerator heat exchanger 510 and an output 521 fluidly coupledto the combustion section of the turbine engine as indicated at 509 maybe included. A gas metering valve 522 may be in fluid communication withthe gas supply line 508. More specifically, the gas metering valve 522may be disposed fluidically in series with the second side 514 of theregenerator heat exchanger 510 and may modulate the flow rate of thegaseous fuel through the second side 514 of the regenerator heatexchanger 510.

During operation, the liquid fuel may flow through the liquid meteringvalve 516, which may modulate the flow rate of the liquid fuel. Theliquid fuel may flow through a first side 512 of the regenerator heatexchanger 510, where it may absorb heat from the second side 514 of theregenerator heat exchanger 510. The liquid fuel may then flow through avaporizer heat exchanger 504, where it may undergo a phase change (e.g.,boiling) to become a gas. The gaseous fuel may flow through the secondside 514 of the regenerator heat exchanger 510, where it may give upheat to the liquid fuel flowing through the first side 512 of theregenerator heat exchanger 510, and/or may flow through a bypass 518around the regenerator heat exchanger 510. The gaseous fuel from thesecond side 514 of the regenerator heat exchanger 510 and/or the bypass518 may be supplied to the combustor 90 of the turbine engine 101 asindicated at 509.

The gas metering valve 522 may control the flow rate of gas through thesecond side 514 of the regenerator heat exchanger 510. Generally, if itis desired to lower the temperature of the gaseous fuel flowing to thecombustor 90, the gas metering valve 522 may be at least partiallyopened to allow more flow through the second side 514 of the regeneratorheat exchanger 510. If it is desired to raise the temperature of thegaseous fuel flowing to the combustor 90, the gas metering valve 522 maybe at least partially shut to allow less flow through the second side514 of the regenerator heat exchanger 510.

It will be understood that any of the above embodiments may be utilizedin a dual fuel aircraft system for a turbine engine that includes afirst fuel a first fuel system for controlling the flow of a first fuelfrom a first fuel tank to the turbine engine and a second fuel systemfor controlling the flow of liquid natural gas to the turbine engine.

The above described embodiments allow for fluid temperature control insingle and dual fuel engines. The above described embodiments providetemperature control for fuels that are to be vaporized and/or need to bebrought to a desired temperature before reaching the combustion sectionof a turbine engine. The above described embodiments provide that aregenerator heat exchanger may be arranged to transfer heat from arelatively warm gaseous fuel to a relatively cool liquid fuel, a bypassmay be arranged to direct at least some gaseous fuel around theregenerator heat exchanger and a metering valve may be arranged tomodulate flow of gaseous fuel through the regenerator heat exchangerand/or through the bypass.

To the extent not already described, the different features andstructures of the various embodiments may be used in combination witheach other as desired. That one feature may not be illustrated in all ofthe embodiments is not meant to be construed that it may not be, but isdone for brevity of description. Thus, the various features of thedifferent embodiments may be mixed and matched as desired to form newembodiments, whether or not the new embodiments are expressly described.All combinations or permutations of features described herein arecovered by this disclosure.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they have structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A turbine engine assembly, comprising: a turbinecore, comprising: a compressor section; a combustion section; and aturbine section, which are all axially aligned; and a cryogenic fuelsystem, comprising: a cryogenic fuel reservoir; a vaporizer heatexchanger; a liquid supply line operably coupling the cryogenic fuelreservoir to an input of the vaporizer heat exchanger; a gas supply lineoperably coupling an output of the vaporizer heat exchanger to thecombustion section; and a second heat exchanger thermally connecting theliquid supply line and the gas supply line to transfer heat from the gassupply line to the liquid supply line.
 2. The engine assembly of claim 1wherein the second heat exchanger is a regenerator heat exchanger. 3.The engine assembly of claim 2 wherein the regenerator heat exchangercomprises a first side fluidly coupled between the fuel reservoir andthe input of the vaporizer heat exchanger.
 4. The engine assembly ofclaim 3 wherein the regenerator heat exchanger comprises a second sidefluidly coupled between the output of the vaporizer heat exchanger andthe combustion section of the turbine engine assembly.
 5. The engineassembly of claim 4, further comprising a bypass line having an inputfluidly located between the output of the vaporizer heat exchanger andthe second side of the regenerator heat exchanger and an output fluidlycoupled to the combustion section of the turbine engine assembly.
 6. Theengine assembly of claim 1, further comprising a liquid metering valvein fluid communication with the liquid supply line and controlling aflow rate of liquid fuel.
 7. The engine assembly of claim 1, wherein thecryogenic fuel is Liquefied Natural Gas (LNG).
 8. The engine assembly ofclaim 2, further comprising a liquid metering valve in fluidcommunication with the liquid supply line and controlling a flow rate ofliquid fuel.
 9. The engine assembly of claim 2, wherein the cryogenicfuel is Liquefied Natural Gas (LNG).
 10. The engine assembly of claim 3,wherein the cryogenic fuel is Liquefied Natural Gas (LNG).
 11. Theengine assembly of claim 3, further comprising a liquid metering valvein fluid communication with the liquid supply line and controlling aflow rate of liquid fuel.
 12. The engine assembly of claim 4, furthercomprising a gas metering valve in fluid communication with the gassupply line and controlling a flow rate of gas through the second sideof the regenerator heat exchanger.
 13. The engine assembly of claim 4,further comprising a liquid metering valve in fluid communication withthe liquid supply line and controlling a flow rate of liquid fuel.
 14. Adual fuel aircraft system for an aircraft turbine engine having acombustion section, comprising: a first fuel system for controlling aflow of a first fuel from a first fuel tank to the aircraft turbineengine; and a second fuel system for controlling a flow of cryogenicfuel to the aircraft turbine engine, comprising: a cryogenic fuelreservoir; a vaporizer heat exchanger; a liquid supply line operablycoupling the cryogenic fuel reservoir to an input of the vaporizer heatexchanger; a gas supply line operably coupling an output of thevaporizer heat exchanger to the combustion section; and a second heatexchanger thermally connecting the liquid supply line and the gas supplyline to transfer heat from the gas supply line to the liquid supplyline.
 15. The dual fuel aircraft system of claim 14 wherein the secondheat exchanger is a regenerator heat exchanger.
 16. The dual fuelaircraft system of claim 15 wherein the regenerator heat exchangercomprises a first side fluidly coupled between the fuel reservoir andthe input of the vaporizer heat exchanger.
 17. The dual fuel aircraftsystem of claim 16 wherein the regenerator heat exchanger comprises asecond side fluidly coupled between the output of the vaporizer heatexchanger and the combustion section of the aircraft turbine engine. 18.The dual fuel aircraft system of claim 17, further comprising a bypassline having an input fluidly located between the output of the vaporizerheat exchanger and the second side of the regenerator heat exchanger andan output fluidly coupled to the combustion section of the aircraftturbine engine.
 19. The dual fuel aircraft system of claim 18, furthercomprising a gas metering valve in fluid communication with the gassupply line and controlling a flow rate of gas through the second sideof the regenerator heat exchanger.
 20. The dual fuel aircraft system ofclaim 14, further comprising a liquid metering valve in fluidcommunication with the liquid supply line and controlling the flow rateof liquid fuel.